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Nasa Tm X-52394 Memorandum NASA TECHNICAL NASA TM X-52394 MEMORANDUM EXPLORING IN AEROSPACE ROCKETRY 7. LIQUID-PROPELLANT ROCKET SYSTEMS by E. William Conrad Lewis Research Center Cleveland, Ohio Presented to Lewis Aerospace Explorers Cleveland, Ohio 1966-67 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION - WASHING , EXPLORING IN AEROSPACE ROCKETRY 7. LIQUID-PROPELLANT ROCKET SYSTEMS E. William Conrad Presented to Lewis Aerospace Explorers Cleveland, Ohio 1966-67 Advisor, James F. Connors NATIONAL AERONAUTICS AND SPACE ADMINISTRATION NASA Technical Chapter Memorandum 1 AEROSPACE ENVIRONMENT John C. Evvard ............................ X-52388 2 PROPULSION FUNDAMENTALS James F. Connors .......................... X-52389 3 CALCULATION OF ROCKET VERTICAL-FLIGHT PERFORMANCE John C. Evvard ............................ X-52390 4 THERMODYNAMICS Marshall C. Burrows ........................ X-52391 5 MATERIALS William D. Klopp ........................... X-52392 6 SOLID-PROPELLANT ROCKET SYSTEMS Joseph F. McBride .......................... X-52393 7 LIQUID-PROPELLANT ROCKET SYSTEMS E. William Conrad .......................... X-52394 8 ZERO-GRAVITY EFFECTS William J. Masica .......................... X-52395 9 ROCKET TRAJECTORIES, DRAG, AND STABILITY Roger W. Luidens .......................... X-52396 10 SPACE MISSIONS Richard J. Weber. .......................... X-52397 11 LAUNCH VEHICLES Arthur V. Zimmerman ........................ X-52398 12 INERTIAL GUIDANCE SYSTEMS Daniel J. Shramo ........................... X-52399 13 TRACKING John L. Pollack. ........................... X-52400 14 ROCKETLAUNCHPHOTOGRAPHY William A. Bowles .......................... X-52401 15 ROCKET MEASUREMENTS AND INSTRUMENTATION Clarence (1. Gettelman ........................ X-52402 16 ELEMENTS OF COMPUTERS Robert L. Miller ............... ........... X-52403 17 ROCKET TESTING AND EVALUATION IN GROUND FACILITIES John H. Povolny. ........................... X-52404 18 LAUNCH OPERATIONS Maynard I. Weston .......................... X-52405 19 NUCLEAR ROCKETS A. F. Lietzke. ............................ X-52406 20 ELECTRIC PROPULSION Harold Kaufman. ........................... X-52407 21 BIOMEDICAL ENGINEERING Kirby W. Hiller. ........................... X-52408 iii 7. LIQUID-PROPELLANT ROCKET SYSTEMS E. William Conrad* Liquid-propellant rockets may be classified as monopropellant, bipropellant, or tri- propellant. In a monopropellant rocket, a propellant, such as hydrazine, is passed through a catalyst to promote a reaction which produces heat from the decomposition of the propellant. A bipropellant rocket burns two chemical materials, a fuel and an oxi- dizer, together. In a tripropellant rocket, three different chemical species, such as hy- drogen, oxygen, and beryllium, are mixed in the combustion chamber and are burned to- gether. These tripropellant rockets have great potential, but they are not yet in actual use because they present many developmental problems. The following discussion will be restricted to bipropellant rockets because this type is used for the bulk of our present space activities . ROCKET ENGINE A simple, liquid-propellant rocket engine is shown in figure 7-1. The principalcom- ponents of this engine are the injector, the combustion chamber, and the exhaust nozzle. rExhaust nozzle \ Combustion 'Figure 7-1. - Bipropellant liquid rocket engine. *Chief, Chemical Rocket Evaluation Branch. The propellants enter the combustion chamber through the injector. In the combustion chamber the propellants mix and are ignited. Some propellants, such as the oxygen- kerosene combinations used in the Atlas launch vehicle, are ignited by means of a spark plug. Other propellants, such as the nitrogen tetroxide - hydrazine combination used in the advanced Titan launch vehicle, are hypergolic; that is, when the two propellants are mixed, they ignite spontaneously. When the propellants burn, they produce very hot gases, The high temperature, in turn, raises the pressure of these gases in the combus- tion chamber. The increased pressure causes the gases to be discharged through the ex- haust nozzle. As these gases pass through the exhaust nozzle, they are accelerated and expanded. The area reduction at the nozzle accelerates the gas to sonic velocity at the throat. Then, in the diverging portion of the nozzle, the gases are expanded and accel- erated to supersonic velocities. (This flow process is discussed in chapter 2. ) Fuel 1 njector The design of the injector is of great importance because the propellants must be introduced into the combustion chamber in such a way that they will mix properly. The objectives of the mixing process are to attain fine atomization of the propellants, rapid evaporation and reaction of the propellants as close as possible to the injector face, and a uniform mixture ratio throughout the combustion chamber, The ultimate goal is to have each molecule of fuel meet an appropriate number of oxidizer molecules and be com- pletely consumed in the combustion process. A detailed discussion of the fundamental processes of combustion is presented in chapter 4. Injectors of many types are used to accomplish the desired mixing of the propellants. Some of the most commonly used injectors are shown in figure 7-2. The double impinging stream injector, shown in figure 7-2(a), is a relatively common design. Fuel and oxi- dizer are supplied to the combustion chamber through alternate manifolds, so that each fuel stream impinges on an oxidizer stream. This impingement shatters the streams into ligaments, which, in turn, break up into droplets, Finally, the droplets evaporate and burn. The triple impinging stream injector (fig, 7-2(b)) is also very common and highly efficient. With this design, two streams of one propellant impinge on a stream of the other propellant at a common point. Figure 7-2(c) shows the self-impinging pattern, in which two streams of the same propellant impinge on each other and shatter to produce a fine, fan-shaped, misty spray. Alternate manifolds in the injector produce fans of fuel mist and of oxidizer mist. These fans mix and burn along their intersections. The shower-head stream injector, shown in figure 7-2(d), was very common in the early days of rocketry. With this design, streams of each propellant are simply injected parallel to one another. The efficiency of this system is, in general, relatively poor. Too much of 2 ,-Impingement boints -Face of injector (a) Double impinging stream. (b) Triple impinging stream. -Typical Fuel impingement ,Typical straight manifolds( point propel la nt stream Oxidizer manifolds( (c) Self-impinging stream. (d) Shower-head stream. ,-Convergent Conical splash plate - cone spray (e) Spray injection. (f) Splash plate. ,-Premix chamber Fuel / Oxidizer swirl volute LFuel inlet (h) Premixing type. (g) Nonimpinging stream. Figure 7-2. - Several injector types. each propellant goes out the exhaust nozzle without mixing and reacting with the other + propellant. Because of this low efficiency, very few current engines use shower-head injectors. Also shown in figure 7-2 are some rather unusual injector designs. The spray injec- tor (fig, 7-2(e)) produces a cone of oxidizer and impinging streams of fuel. Figure 7-2(f) shows a splash-plate injector, in which streams of fuel and oxidizer impinge at a point on the splash plate. This impingement produces sprays that eddy around the splash plate and promote further mixing of the propellants. The nonimpinging stream injector, shown in figure 7-2(g), has a precombustion chamber in the form of a cup sunk into the injector face. Many streams of both propellants are injected into this precombustion chamber to produce a rather violent mixing. The premixing injector (fig. 7-2(h)) has a premixing chamber into which the two propellants are injected tangentially to produce a swirling mixing action, There is a new and relatively efficient injector which uses a quadruple impinging stream pattern. With this design, two streams of each propellant impinge at a common point, This injector is particularly effective for use with storable propellants. The concentric tube injector, shown in figure 7-3, is probably the optimum design for hydrogen-oxygen propellant combinations The oxidizer enters the oxidizer cavity through the center pipe, then flows outward throughout this cavity, and enters the combus- tion chamber through the hollow oxidizer tubes. The fuel enters the fuel cavity, which is just under the injector face, and thence it flows into the combustion chamber through the annuli which surround the oxidizer tubes. I Figure 7-3. - Cross section of concentric-tube injector. 4 Corn bustion Instability All these injection techniques create a combustion zone which has a great deal of energy contained in it; this concentration of energy can cause severe problems. One great difficulty in developing new rocket engines is combustion instability, particularly the variety at high frequency which we call "screech" or "screaming. ?' This phenom- enon has plagued propulsion people since afterburners were developed in the late 1940's and has continued through ramjets and into the rocket field. Screech can increase heat transfer by a factor of as much as 10 and thus is extremely destructive. As an example, in figure 7-4 is shown an injector face which experienced screech for only 0.4 second. Why screech happens is not yet fully understood, but there has been, and there still is, a great effort aimed at trying to solve the combustion instability problem. What apparently happens in the engine is that pressure
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